US2025314208A1PendingUtilityA1
High power epicyclic gearbox and operation thereof
Est. expiryDec 5, 2039(~13.3 yrs left)· nominal 20-yr term from priority
Inventors:Mark Spruce
F05D 2200/14F05D 2220/36F05D 2220/323F05D 2260/40311F02K 3/06Y02T50/60F05D 2300/501F05D 2260/96B64D 27/10F02C 3/04F02C 7/36F16H 57/025F16H 57/023F16H 57/08F16H 48/10F16H 1/32
89
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Claims
Abstract
A gas turbine engine for an aircraft including: an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan including a plurality of fan blades; a gearbox that can receive an input from the core shaft, and can output drive to a fan shaft via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft; and a fan shaft mounting structure arranged to mount the fan shaft within the engine, the fan shaft mounting structure including at least two supporting bearings connected to the fan shaft.
Claims
exact text as granted — not AI-modified1 . A gas turbine engine for an aircraft comprising:
an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that is configured to:
receive an input from the core shaft, and
output drive to a fan shaft via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft; and
a fan shaft mounting structure arranged to mount the fan shaft within the gas turbine engine, the fan shaft mounting structure comprising at least two supporting bearings connected to the fan shaft, wherein: the output of the gearbox is at a gearbox output position and the input to the fan is at a fan input position; a first bearing separation distance (d 1 ) is defined as an axial distance between the input to the fan and a closest bearing of the at least two supporting bearings in a rearward direction from the fan; both of the at least two supporting bearings are located at positions forward of the gearbox; a bearing axial separation (d 3 ) is defined as an axial distance between the closest bearing of the at least two supporting bearings in the rearward direction from the fan and a closest bearing of the at least two supporting bearings in a forward direction from the gearbox; a bearing separation ratio defined as:
the
first
bearing
separation
distance
(
d
1
)
the
bearing
axial
separation
(
d
3
)
is greater than 4.1×10 −1 ; and
a product of:
(a radial bending stiffness of the fan shaft at the input to the fan)×(a radial bending stiffness of the fan shaft at the output of the gearbox)
is greater than or equal to 1.2×10 13 (N/m) 2 .
2 . The gas turbine engine according to claim 1 , wherein the bearing separation ratio is in a range of 4.1×10 −1 to 8.3×10 1 .
3 . The gas turbine engine according to claim 1 , wherein the bearing separation ratio is in a range of 4.1×10 −1 to 6.0×10 1 .
4 . The gas turbine engine according to claim 1 , wherein the bearing separation ratio is in a range of 6.0×10 −1 to 8.3×10 −1 .
5 . The gas turbine engine according to claim 1 , wherein at least one of the following is satisfied:
a tilt stiffness of the fan shaft at the input to the fan is in a range from 5.00×10 5 Nm/rad to 7.00×10 8 Nm/rad; and a tilt stiffness of the fan shaft at the output of the gearbox is in a range from 7.00×10 4 Nm/rad to 7.00×10 7 Nm/rad.
6 . The gas turbine according to claim 1 , wherein the gearbox is in a planetary configuration and the output of the gearbox is a gearbox output position at an interface between the fan shaft and a planet carrier.
7 . The gas turbine engine according to claim 1 , wherein the bearing axial separation (d 3 ) is in a range from 0.18 m to 0.32 m.
8 . The gas turbine engine according to claim 1 , wherein:
each fan blade comprises a carbon-fibre based body; and the material from which the fan shaft is made has a Young's modulus in the range from 105 to 215 GPa.
9 . The gas turbine engine according to claim 1 , wherein at least one of the following is satisfied:
the first bearing separation distance (d 1 ) is in a range from 0.12 m to 0.40 m; a second bearing separation distance (d 2 ) is defined as an axial distance between the output of the gearbox and the closest bearing of the at least two supporting bearings in the forward direction from the gearbox, the second bearing separation distance (d 2 ) being in a range from 0.15 m to 0.45 m; and an axial distance between the fan input position and the gearbox output position (d 4 ) is in a range from 0.43 m to 0.95 m.
10 . A gas turbine engine for an aircraft comprising:
an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that is configured to:
receive an input from the core shaft, and
output drive to a fan shaft via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft; and
a fan shaft mounting structure arranged to mount the fan shaft within the gas turbine engine, the fan shaft mounting structure comprising at least two supporting bearings connected to the fan shaft, wherein: the output of the gearbox is at a gearbox output position and the input to the fan is at a fan input position; a first bearing separation distance (d 1 ) is defined as an axial distance between the input to the fan and a closest bearing of the at least two supporting bearings in a rearward direction from the fan; both of the at least two supporting bearings are located at positions forward of the gearbox; a bearing axial separation (d 3 ) is defined as an axial distance between the closest bearing of the at least two supporting bearings in the rearward direction from the fan and a closest bearing of the at least two supporting bearings in a forward direction from the gearbox; a bearing separation ratio defined as:
the
first
bearing
separation
distance
(
d
1
)
the
bearing
axial
separation
(
d
3
)
is greater than 4.1×10 −1 ; and
an overall pressure ratio defined as the ratio of a stagnation pressure upstream of the fan to a stagnation pressure at the exit of a highest pressure section of the compressor is in a range of 45 to 60 at cruise conditions.
11 . The gas turbine engine according to claim 10 , wherein the bearing separation ratio is in a range of 4.1×10 −1 to 6.0×10 −1 .
12 . The gas turbine engine according to claim 10 , wherein the bearing separation ratio is in a range of 6.0×10 −1 to 8.3×10 −1 .
13 . The gas turbine engine according to claim 10 , wherein a further bearing separation ratio of:
the
first
bearing
separation
distance
(
d
1
)
the
axial
distance
between
the
fan
input
position
and
the
gearbox
output
position
(
d
4
)
is greater than or equal to 1.6×10 −1 .
14 . The gas turbine engine according to claim 13 , wherein the further bearing separation ratio is in a range from 1.6×10 −1 to 2.2×10 −1 .
15 . A gas turbine engine for an aircraft comprising:
an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that is configured to:
receive an input from the core shaft, and
output drive to a fan shaft via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft; and
a fan shaft mounting structure arranged to mount the fan shaft within the gas turbine engine, the fan shaft mounting structure comprising at least two supporting bearings connected to the fan shaft, wherein: the output of the gearbox is at a gearbox output position and the input to the fan is at a fan input position; a first bearing separation distance (d 1 ) is defined as an axial distance between the input to the fan and a closest bearing of the at least two supporting bearings in a rearward direction from the fan; both of the at least two supporting bearings are located at positions forward of the gearbox; a bearing axial separation (d 3 ) is defined as an axial distance between the closest bearing of the at least two supporting bearings in the rearward direction from the fan and a closest bearing of the at least two supporting bearings in a forward direction from the gearbox; a bearing separation ratio defined as:
the
first
bearing
separation
distance
(
d
1
)
the
bearing
axial
separation
(
d
3
)
is greater than 4.1×10 −1 ; and either one or both of the following is satisfied:
(i) a radial bending stiffness of the fan shaft mounting structure is greater than or equal to 7.00×10 8 N/m; or
(ii) a tilt stiffness of the fan shaft mounting structure is greater than or equal to 1.5×10 7 Nm/rad.
16 . The gas turbine engine according to claim 15 , wherein the bearing separation ratio is in a range of 4.1×10 1 to 6.0×10 1 .
17 . The gas turbine engine according to claim 15 , wherein the bearing separation ratio is in a range of 6.0×10 −1 to 8.3×10 −1 .
18 . The gas turbine engine according to claim 15 , wherein the radial bending stiffness of the fan shaft mounting structure is in a range from 1.25×10 9 N/m to 2.0×10 11 N/m.
19 . The gas turbine engine according to claim 15 , wherein the tilt stiffness of the fan shaft mounting structure is in a range from 2.1×10 7 Nm/rad to 1×10 10 Nm/rad.
20 . The gas turbine engine according to claim 15 , wherein at least two of the following are satisfied:
the first bearing separation distance (d 1 ) is in a range from 0.12 m to 0.40 m; a second bearing separation distance (d 2 ) is defined as an axial distance between the output of the gearbox and the closest bearing of the at least two supporting bearings in the forward direction from the gearbox, the second bearing separation distance (d 2 ) being in a range from 0.15 m to 0.45 m; the bearing axial separation (d 3 ) is in a range from 0.18 m to 0.32 m; and an axial distance between the fan input position and the gearbox output position (d 4 ) is in a range from 0.43 m to 0.95 m.Join the waitlist — get patent alerts
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